GB1458066A - Gas turbine combustion chambers - Google Patents

Abstract

1458066 Gas turbine combustion chamber SOC NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION 12 March 1974 [13 March 1973] 10941/74 Heading F1L To reduce the production of pollutants in a gas turbine combustion chamber 1 for jet aircraft, the chamber 1 is formed as two coaxial annular chambers receiving air from a compressor 3 at the same or different pressures, an outer chamber 6 of conventional flame tube type having fuel injected continuously at a slow rate at 24 and combustion air supplied to ensure a substantially stoichiometric mixture with the slow rate of fuel feed for slow running of the engine, and an inner chamber 8 having fast fuel feed at 40 in an upstream air inlet 22a of convergent/divergent form for use with the slow fuel feed in chamber 1 during take-off. At cruising, fuel supply is cut-off from injectors 40 and fed either as shown into a secondary combustion zone 29 in chamber 6 via injectors 31 or, into the inner chamber 8 via injectors (54), Fig. 6, (not shown), formed integral with the downstream flame stabilizers 45. The downstream end of chamber 6 is partially closed off by a plate 37 formed with apertures 38 to form jets for mixing the combustion products from chambers 6 and 8 in the downstream mixing zone 39 prior to entry to the turbine 4. Again Figs. 5 and 5a, (not shown), mixing may be effected by a corrugated baffle (53) positioned on the downstream end of the U-shaped wall 11 inter-connecting the two chamber 6 and 8. Communicating ducts 44 between the chambers ensures ignition in chamber 8.